Known turbine engines, including gas/combustion turbine engines and steam turbine engines, incorporate shaft-mounted turbine blades circumferentially circumscribed by a turbine casing or housing. The remainder of this description focuses on applications within combustion or gas turbine technical application and environment, though exemplary embodiments described herein are applicable to steam turbine engines. In a gas/combustion turbine engine hot combustion gasses flow in a combustion path that initiates within a combustor and are directed through a generally tubular transition into a turbine section. A forward or Row 1 vane directs the combustion gasses past successive alternating rows of turbine blades and vanes. Hot combustion gas striking the turbine blades cause blade rotation, thereby converting thermal energy within the hot gasses to mechanical work, which is available for powering rotating machinery, such as an electrical generator.
Engine internal components within the hot combustion gas path are exposed to combustion temperatures on the order of 900 degrees Celsius (1600 degrees Fahrenheit). The engine internal components within the combustion path, such as for example combustion section transitions, vanes and blades are often constructed of high temperature resistant superalloys. Blades and vanes often include cooling passages terminating in cooling holes on component outer surface, for passage of coolant fluid into the combustion path.
Turbine engine internal components often incorporate a thermal barrier coat or coating (TBC) of metal-ceramic material that is applied directly to the external surface of the component substrate surface or over an intermediate metallic bond coat (BC) that was previously applied to the substrate surface. The TBC provides an insulating layer over the component substrate, which reduces the substrate temperature. Combination of TBC application along with cooling passages in the component further lowers the substrate temperature.
Due to differences in thermal expansion, fracture toughness and elastic modulus—among other things—between typical metal-ceramic TBC materials and typical superalloy materials used to manufacture the aforementioned exemplary turbine components, there is potential risk of cracking the TBC layer as well as TBC/turbine component adhesion loss at the interface of the dissimilar materials. The cracks and/or adhesion loss/delamination negatively impact the TBC layer structural integrity and potentially lead to its spallation, i.e., separation of the insulative material from the turbine component. For example, vertical cracks developing within the TBC layer can propagate to the TBC/substrate interface, and then spread horizontally. Similarly, horizontally oriented cracks can originate within the TBC layer or proximal the TBC/substrate interface. Such cracking loss of TBC structural integrity can lead to further, premature damage to the underlying component substrate. When the TBC layer breaks away from underlying substrate the latter loses its protective thermal layer coating. During continued operation of the turbine engine it is possible over time that the hot combustion gasses will erode or otherwise damage the exposed component substrate surface, potentially reducing engine operation service life. Potential spallation risk increases with successive powering on/off cycles as the engine is brought on line to generate electrical power in response to electric grid increased load demands and idling down as grid load demand decreases. In order to manage the TBC spallation risk and other engine operational maintenance needs, combustion turbine engines are often taken out of service for inspection and maintenance after a defined number of powering on/off thermal cycles.
In addition to thermal or vibration stress crack susceptibility, the TBC layer on engine components is also susceptible to foreign object damage (FOD) as contaminant particles within the hot combustion gasses strike the relatively brittle TBC material. A foreign object impact can crack the TBC surface, ultimately causing spallation loss of surface integrity that is analogous to a road pothole. Once foreign object impact spalls of a portion of the TBC layer, the remaining TBC material is susceptible to structural crack propagation and/or further spatting of the insulative layer. In addition to environmental damage of the TBC layer by foreign objects, contaminants in the combustion gasses, such as calcium, magnesium, aluminum and silicon (often referred to as “CMAS”) can adhere to or react with the TBC layer, increasing the probability of TBC spallation and exposing the underlying bond coat.
Past attempts to enhance TBC layer structural integrity and affixation to underlying turbine component substrates have included development of stronger TBC materials better able to resist thermal cracking or FOD, but with tradeoffs in reduced thermal resistivity or increased material cost. Generally the relatively stronger, less brittle potential materials for TBC application have had lower thermal resistivity. Alternatively, as a compromise separately applied multiple layers of TBC materials having different advantageous properties have been applied to turbine component substrates—for example a more brittle or softer TBC material having better insulative properties that is in turn covered by a stronger, lower insulative value TBC material as a tougher “armor” outer coating better able to resist FOD and/or CMAS contaminant adhesion. In order to improve TBC adhesion to the underlying substrate, intermediate metallic bond coat (BC) layers have been applied directly over the substrate. Structural surface properties and/or profile of the substrate or BC interface to the TBC have also been modified from a flat, bare surface. Some known substrate and/or BC surface modifications (e.g., so-called “rough bond coats” or RBCs) have included roughening the surface by ablation or other blasting, thermal spray deposit or the like. In some instances the BC or substrate surface has been photoresist or laser etched to include surface features on the order of a few microns (μm) height and spacing width across the surface planform. Features have been formed directly on the substrate surface of turbine blade tips to mitigate stress experienced in blade tip coatings. Rough bond coats have been thermally sprayed to leave porous surfaces of a few micron-sized features. TBC layers have been applied by locally varying homogeneity of the applied ceramic-metallic material to create pre-weakened zones for attracting crack propagation in controlled directions. For example a weakened zone has been created in the TBC layer corresponding to a known or likely stress concentration zone, so that any cracks developing therein are propagated in a desired direction to minimize overall structural damage to the TBC layer.